The present disclosure relates generally to gas turbine engines, any turbomachinery, and, more specifically, to turbines therein.
In a gas turbine engine air is pressurized in a compressor and mixed with fuel in a combustor for generating hot combustion gases. Turbine stages extract energy from the combustion gases to power the compressor, while also powering an upstream fan in a turbofan aircraft engine application, or powering an external drive shaft for marine and industrial applications.
In aviation applications, a high pressure turbine (HPT) immediately follows the combustor and includes a stationary turbine nozzle which discharges combustion gases into a row of rotating first stage turbine rotor blades extending radially outwardly from a supporting rotor disk. The HPT may include one or more stages of rotor blades and corresponding turbine nozzles. Following the HPT is a low pressure turbine (LPT) which typically includes multiple stages of rotor blades and corresponding turbine nozzles. In alternate applications, such as land based gas turbines for power generation, embodiments may employ a single shaft turbine, with or without a low pressure turbine (LPT).
Each turbine nozzle includes a row of stator vanes having radially outer and inner endwalls in the form of arcuate bands which support the vanes. Correspondingly, the turbine rotor blades include airfoils integrally joined to radially inner endwalls or platforms supported in turn by corresponding dovetails which provide mounting of the individual blades in dovetail slots formed in the perimeter of the supporting rotor disk. An annular shroud surrounds the radially outer tips of the rotor airfoils in each turbine stage.
The stator vanes and rotor blades have corresponding airfoils including generally concave pressure sides and generally convex suction sides extending axially in chord between opposite leading and trailing edges. Adjacent vanes and adjacent blades form corresponding flow passages therebetween bound by the radially inner and outer endwalls.
During operation, combustion gases are discharged from the combustor and flow axially downstream as a core flow through the respective flow passages defined between the stator vanes and rotor blades. In addition, purge air from a purge cavity existing upstream of the airfoil leading edge is discharged as a purge flow that prevents ingesting hot core flow below the main gas path and potentially provides a cooling effect to the platforms and airfoils. The aerodynamic contours of the vanes and blades, and corresponding flow passages therebetween, are precisely configured for maximizing energy extraction from the combustion gases which in turn rotate the rotor from which the blades extend.
The complex three-dimensional (3D) configuration of the vane and blade airfoils is tailored for maximizing efficiency of operation, and varies radially in span along the airfoils as well as axially along the chords of the airfoils between the leading and trailing edges. Accordingly, the velocity and pressure distributions of the combustion gases and purge air over the airfoil surfaces, as well as within the corresponding flow passages, also vary.
Undesirable pressure losses in the combustion gas flowpaths therefore correspond with undesirable reduction in turbine aerodynamics and overall turbine efficiency. For example, the combustion gases enter the corresponding rows of vanes and blades in the flow passages therebetween and are necessarily split at the respective leading edges of the airfoils. In addition, mixing of the purge air flow and the core flow may lead to turbine inefficiency.
The locus of stagnation points of the incident combustion gases extends along the leading edge of each airfoil, and corresponding boundary layers are formed along the pressure and suction sides of each airfoil, as well as along each radially outer and inner endwall which collectively bound the four sides of each flow passage. In the boundary layers, the local velocity of the combustion gases varies from zero along the endwalls and airfoil surfaces to the unrestrained velocity in the combustion gases where the boundary layers terminate.
Turbine losses can occur from a variety of sources, for example, secondary flows, shock loss mechanism and mixing losses. One common source of turbine pressure losses is the formation of horseshoe vortices generated as the combustion gases are split in their travel around the airfoil leading edges. A total pressure gradient is affected in the boundary layer flow at the junction of the leading edge and endwalls of the airfoil. This pressure gradient at the airfoil leading edges forms a pair of counterrotating horseshoe vortices which travel downstream on the opposite sides of each airfoil near the endwall. The two vortices travel aft along the opposite pressure and suction sides of each airfoil and behave differently due to the different pressure and velocity distributions therealong. For example, computational analysis indicates that the suction side vortex migrates away from the endwall toward the airfoil trailing edge and then interacts following the airfoil trailing edge with the pressure side vortex flowing aft thereto.
The interaction of the pressure and suction side vortices occurs near the midspan region of the airfoils and creates total pressure loss and a corresponding reduction in turbine efficiency. These vortices also create turbulence and increase undesirable heating of the endwalls.
Since the horseshoe vortices are formed at the junctions of turbine rotor blades and their integral root platforms, as well at the junctions of nozzle stator vanes and their outer and inner bands, corresponding losses in turbine efficiency are created, as well as additional heating of the corresponding endwall components.
Similarly, cross-passage pressure gradients between the pressure and suction side of the blade give rise to secondary flow structures and vortices that alter the desired aerodynamics of the blade, giving rise to losses in turbine efficiency as well as possible heating of the endwalls and even the blade.
At the leading edges of the turbine blades, and more particularly at a junction of the leading edge and the leading edge purge cavity, secondary flow structures and mixing of a purge flow from the leading edge purge cavity, results in mixing losses. In addition, the secondary flow structures result in mixing of the purge flow with the main core flow, resulting in a trajectory of the purge flow that is remote from the platform. These secondary flow structures result in high heat concentrations in the area where the turbine blade join the blade endwall structure.
Accordingly, it is desired to provide an improved turbine stage for reducing horseshoe and secondary flow vortex affects, as well as increasing aerodynamic loading while controlling heat distribution and efficiency, or improving efficiency and thermal loading, while maintaining aerodynamic loading and/or torque production.